Turbine blade

ABSTRACT

An airfoil comprises one or more internal cooling circuits. The cooling circuit can further comprise a near wall cooling mesh, fluidly coupling a supply passage to a mesh plenum. The mesh plenum can be disposed adjacent to the external surface of the airfoil having a plurality of film holes extending between the mesh plenum and the external surface of the airfoil. The mesh plenum can further comprise a cross-sectional area sized to facilitate machining of the film holes without damage to the interior of the airfoil.

CROSS-REFERENCE TO RELATED APPLICATION

This application is a continuation of U.S. application Ser. No.14/884,075, filed on Oct. 15, 2015, titled “TURBINE BLADE”, which ishereby expressly incorporated herein by reference in its entirety.

BACKGROUND OF THE INVENTION

Turbine engines, and particularly gas or combustion turbine engines, arerotary engines that extract energy from a flow of combusted gasespassing through the engine onto a multitude of rotating turbine blades.Gas turbine engines have been used for land and nautical locomotion andpower generation, but are most commonly used for aeronauticalapplications such as for aircraft, including helicopters. In aircraft,gas turbine engines are used for propulsion of the aircraft. Interrestrial applications, turbine engines are often used for powergeneration.

Gas turbine engines for aircraft are designed to operate at hightemperatures to maximize engine efficiency, so cooling of certain enginecomponents, such as the high pressure turbine and the low pressureturbine, can be beneficial. Typically, cooling is accomplished byducting cooler air from the high and/or low pressure compressors to theengine components that require cooling. Temperatures in the highpressure turbine are around 1000° C. to 2000° C. and the cooling airfrom the compressor is around 500° C. to 700° C. While the compressorair is a high temperature, it is cooler relative to the turbine air, andcan be used to cool the turbine.

Contemporary turbine blades generally include one or more interiorcooling circuits for routing the cooling air through the blade to cooldifferent portions of the blade, and can include dedicated coolingcircuits for cooling different portions of the blade, such as theleading edge, trailing edge and tip of the blade.

BRIEF DESCRIPTION OF THE INVENTION

An airfoil having an outer surface defining a pressure side and asuction side extending axially between a leading edge and a trailingedge and extending radially between a root and a tip. The airfoilfurther comprises a cooling circuit located within the airfoil furthercomprising a radially extending supply passage fluidly coupled to thecooling air inlet passage, a near wall cooling mesh located adjacent toand extending along a portion of the outer surface, a radially extendingopening fluidly coupling the supply passage to the near wall coolingmesh to define a fluid inlet for the near wall cooling mesh, and aplenum fluidly coupled to the near wall cooling mesh to define a fluidoutlet for the near wall cooling mesh, wherein cooling air flows throughthe near wall cooling mesh from the inlet to the outlet and thecross-sectional area of the plenum is greater than the cross-sectionalarea of the cooling mesh in the flow direction.

A blade for a gas turbine engine having a turbine rotor disk. The bladecomprises a dovetail having multiple air inlet passages and configuredto mount to the turbine rotor disk. The blade further comprises anairfoil extending radially from the dovetail and having an outer surfacedefining a pressure side and a suction side extending axially betweenthe leading edge and a trailing edge and extending radially between aroot and a tip, with the root being adjacent the dovetail. The bladefurther comprises a leading edge cooling circuit located within theairfoil and fluidly coupled to a corresponding one of the multiple airinlet passages, a trailing edge cooling circuit located within theairfoil and fluidly coupled to a corresponding one of the multiple airinlet passages, and a mid cooling circuit located within the airfoil andbetween the leading edge cooling circuit and the trailing edge coolingcircuit, the mid cooling circuit comprising: a radially extending supplypassage fluidly coupled to the cooling air inlet passage, a near wallcooling mesh located adjacent to and extending along a portion of theouter surface, a radially extending opening fluidly coupling the supplypassage to the near wall cooling mesh to define a fluid inlet for thenear wall cooling mesh, and a plenum fluidly coupled to the near wallcooling mesh to define a fluid outlet for the near wall cooling mesh.The cooling air flows through the near wall cooling mesh from the inletto the outlet and the cross-sectional area of het plenum is greater thanthe cross-sectional area of the cooling mesh in the flow direction.

An airfoil for a gas turbine engine having a peripheral wall bounding aninterior and defining a pressure side and a suction side, opposite thepressure side, a near wall cooling mesh located adjacent to andextending along a portion of one of the suction side and the pressureside, and a plenum fluidly coupled to the near wall cooling mesh todefine a fluid outlet for the near wall cooling mesh, wherein the plenumhas a larger cross dimension in a chord-wise plan than the near wallcooling mesh.

BRIEF DESCRIPTION OF THE DRAWINGS

In the drawings:

FIG. 1 is a schematic cross-sectional diagram of a gas turbine enginefor an aircraft.

FIG. 2 is a perspective view of an engine component in the form of aturbine blade of the engine of FIG. 1 with cooling air inlet passages.

FIG. 3 is a cross-sectional view of the airfoil of FIG. 2.

FIG. 4 is a diagram of the cross-sectional airfoil of FIG. 3illustrating four cooling circuits.

FIG. 5 is an enlarged view of a mid cooling circuit of FIG. 4illustrating a near wall cooling mesh and plenum.

FIG. 6 is a flow diagram for the cooling circuits of FIG. 4.

FIG. 7 is a diagram of the cross-sectional airfoil of FIG. 3illustrating a near wall cooling mesh and plenum being fed in theopposite direction than that of FIG. 4.

DESCRIPTION OF EMBODIMENTS OF THE INVENTION

The described embodiments of the present invention are directed to aturbine blade, and in particular to cooling a turbine blade. Forpurposes of illustration, the present invention will be described withrespect to a turbine blade for an aircraft gas turbine engine. It willbe understood, however, that the invention is not so limited and canhave general applicability in non-aircraft applications, such as othermobile applications and non-mobile industrial, commercial, andresidential applications. It can also have application to airfoils,other than a blade, in a turbine engine, such as stationary vanes.

It should be appreciated that as described herein the term “span-wise”should be understood as the direction generally extending between a rootand a tip of an airfoil. It should be further understood that asdescribed herein, the term “chord-wise” should be understood as thedirection generally extending arcuately between a leading edge and atrailing edge of an airfoil. Furthermore, “chord-wise” can define a“chord-wise plane” such that a planar area can be defined extendingbetween the leading edge and the trailing edge.

FIG. 1 is a schematic cross-sectional diagram of a gas turbine engine 10for an aircraft. The engine 10 has a generally longitudinally extendingaxis or centerline 12 extending forward 14 to aft 16. The engine 10includes, in downstream serial flow relationship, a fan section 18including a fan 20, a compressor section 22 including a booster or lowpressure (LP) compressor 24 and a high pressure (HP) compressor 26, acombustion section 28 including a combustor 30, a turbine section 32including a HP turbine 34, and a LP turbine 36, and an exhaust section38.

The fan section 18 includes a fan casing 40 surrounding the fan 20. Thefan 20 includes a plurality of airfoils in the form of fan blades 42disposed radially about the centerline 12. The HP compressor 26, thecombustor 30, and the HP turbine 34 form a core 44 of the engine 10,which generates combustion gases. The core 44 is surrounded by corecasing 46 which can be coupled with the fan casing 40.

A HP shaft or spool 48 disposed coaxially about the centerline 12 of theengine 10 drivingly connects the HP turbine 34 to the HP compressor 26.A LP shaft or spool 50, which is disposed coaxially about the centerline12 of the engine 10 within the larger diameter annular HP spool 48,drivingly connects the LP turbine 36 to the LP compressor 24 and fan 20.

The LP compressor 24 and the HP compressor 26 respectively include aplurality of compressor stages 52, 54, in which a set of rotatingairfoils in the form of compressor blades 56, 58 that rotate relative toa corresponding set of static airfoils in the form of compressor vanes60, 62 (also called a nozzle) to compress or pressurize the stream offluid passing through the stage. In a single compressor stage 52, 54,multiple compressor blades 56, 58 may be provided in a ring and mayextend radially outwardly relative to the centerline 12, from a bladeplatform to a blade tip, while the corresponding static compressor vanes60, 62 are positioned downstream of and adjacent to the rotating blades56, 58. It is noted that the number of blades, vanes, and compressorstages shown in FIG. 1 were selected for illustrative purposes only, andthat other numbers are possible.

The HP turbine 34 and the LP turbine 36 respectively include a pluralityof turbine stages 64, 66, in which a set rotating airfoils in the formof turbine blades 68, 70 that are rotated relative to a correspondingset of static airfoils in the form of turbine vanes 72, 74 (also calleda nozzle) to extract energy from the stream of fluid passing through thestage. In a single turbine stage 64, 66, multiple turbine blades 68, 70may be provided in a ring and may extend radially outwardly relative tothe centerline 12, from a blade platform to a blade tip, while thecorresponding static turbine vanes 72, 74 are positioned upstream of andadjacent to the rotating blades 68, 70. It is noted that the number ofblades, vanes, and turbine stages shown in FIG. 1 were selected forillustrative purposes only, and that other numbers are possible.

In operation, the rotating fan 20 supplies ambient air to the LPcompressor 24, which then supplies pressurized ambient air to the HPcompressor 26, which further pressurizes the ambient air. Thepressurized air from the HP compressor 26 is mixed with fuel in thecombustor 30 and ignited, thereby generating combustion gases. Some workis extracted from these gases by the HP turbine 34, which drives the HPcompressor 26. The combustion gases are discharged into the LP turbine36, which extracts additional work to drive the LP compressor 24, andthe exhaust gas is ultimately discharged from the engine 10 via theexhaust section 38. The driving of the LP turbine 36 drives the LP spool50 to rotate the fan 20 and the LP compressor 24.

Some of the ambient air supplied by the fan 20 may bypass the enginecore 44 and be used for cooling of portions, especially hot portions, ofthe engine 10, and/or used to cool or power other aspects of theaircraft. In the context of a turbine engine, the hot portions of theengine are normally downstream of the combustor 30, especially theturbine section 32, with the HP turbine 34 being the hottest portion asit is directly downstream of the combustion section 28. Other sources ofcooling fluid may be, but is not limited to, fluid discharged from theLP compressor 24 or the HP compressor 26.

FIG. 2 is a perspective view of an engine component in the form of oneof the turbine blades 68 of the engine 10 from FIG. 1. The turbine blade68 includes a dovetail 76 and an airfoil 78. The airfoil 78 extends froma tip 80 to a root 82. The dovetail 76 further includes a platform 84,integral with the airfoil 78 at the root 82, which helps to radiallycontain the turbine air flow. The dovetail 76 can be configured to mountto a turbine rotor disk on the engine 10. The dovetail 76 comprises atleast one inlet passage, exemplarily shown as a first inlet passage 88,a second inlet passage 90, and a third inlet passage 92, each extendingthrough the dovetail 76 to provide internal fluid communication with theairfoil 78 at a passage outlet 94. It should be appreciated that thedovetail 76 is shown in cross-section, such that the inlet passages 88,90, 92 are housed within the body of the dovetail 76.

Turning to FIG. 3, the airfoil 78, shown in cross-section, has aninterior 96, defined by a concave-shaped pressure sidewall 98, defininga pressure side, and a convex-shaped suction sidewall 100, defining asuction side, which are joined together to define the airfoil shape witha leading edge 102 and a trailing edge 104. The airfoil 78 rotates in adirection such that the pressure sidewall 98 follows the suctionsidewall 100. Thus, as shown in FIG. 3, the airfoil 78 would rotateupward toward the top of the page.

The airfoil 78 comprises a plurality of internal passages which can bearranged to form cooling circuits dedicated to cool a particular portionof the blade 68. The passages and the corresponding cooling circuits areillustrated in FIG. 4, which is a cross-sectional view of the airfoil78. It should be appreciated that the respective geometries of eachindividual passage within the airfoil 78 as shown is exemplary, eachdepicting one or more elements of a cooling circuit and should not limitthe cooling circuits to the geometries, dimensions, or positions asshown.

As illustrated, the airfoil 78 is provided with three cooling circuitscomprising a leading edge cooling circuit 120, a mid cooling circuit122, and a trailing edge cooling circuit 124, which are supplied coolingair via the inlet passages 88, 90, 92, respectively. The trailing edgecooling circuit 124, as illustrated, further comprises a first coolingcircuit 124 a and a second cooling circuit 124 b, commonly fed from thethird cooling passage 92.

The cooling circuits can be defined by one or more passages extendingradially within the airfoil 78. It should be appreciated that thepassages can comprise one or more film holes which can provide fluidcommunication between the particular passage and the external surface ofthe airfoil 78, providing a film of cooling fluid along the externalsurface of the airfoil 78.

Looking in greater detail at each of the cooling circuits, the leadingedge cooling circuit 120 can comprise a supply passage 130, near wallcooling circuit 132, and leading edge passage 144. The supply passage130 extends from the root 82 toward the tip 80, being fluidly coupled tothe first inlet passage 88 at the outlet 94 to supply the cooling air tothe near wall cooling circuit 132 and leading edge passage 144.

The near wall cooling circuit 132 is exemplarily illustrated between thesupply passage 130 and the suction sidewall 100, while being adjacent tothe suction sidewall 100. In this configuration, the near wall coolingcircuit 132 cools the wall portion of the airfoil 78 along the suctionsidewall 100. Alternatively, the near wall cooling circuit 132 can bepositioned adjacent to the pressure sidewall 98, between the pressuresidewall 98 and the supply passage 130. As such, the supply passage 130can alternatively be positioned adjacent to the suction sidewall 100.

The near wall cooling circuit 132 comprises a plenum passage 136,extending from the tip 80 to the root 82, and can have a plurality ofpins or pin banks 138 disposed within the plenum passage 136. The nearwall cooling circuit 132 further comprises at least one return passage140, illustrated in FIG. 4 as two return passages 140 located onopposite chord-wise ends of the plenum passage 136. The return passages140 fluidly couple to the plenum passage 136 near the root 82, andextend from the root 82 to the tip 80.

The leading edge cooling circuit 120 further comprises a leading edgepassage 144, being located adjacent to the leading edge 102 andextending from the root 82 to the tip 80. The leading edge passage 144is in fluid communication with the supply passage 130 through a leadcircuit channel 146 defined in a wall 148 between the supply passage 130and the leading edge passage 144. The lead circuit channel 146 cancomprise multiple discrete inlets, such as impingement openings. Theleading edge passage 144 can further be adjacent to one return passage140 of the near wall cooling circuit 132.

The trailing edge cooling circuit 124, comprising the first coolingcircuit 124 a and the second cooling circuit 124 b, can be commonly fedfrom the third inlet passage 92. The third inlet passage 92 can feedboth the first cooling circuit 124 a and the second cooling circuit 124b where the third inlet passage 92 can split from one inlet into twodistinct inlets within the dovetail 76. Alternatively, a fourth inletpassage (not shown) can be used to feed one of the first cooling circuit124 a or the second cooling circuit 124 b while the third inlet passage92 feeds the other. The second cooling circuit 124 b is disposedadjacent to the trailing edge 104 and the first cooling circuit 124 a isdisposed chord-wise between the second cooling circuit 124 b and the midcooling circuit 122.

The first cooling circuit 124 a comprises a supply passage 150, being influid communication with the third inlet passage 92, and extending fromthe root 82 to the tip 80. The supply passage 150 fluidly couples to areturn passage 152 extending from the tip 80 to the root 82, whichfluidly couples to an outlet passage 154 extending from the root 82 tothe tip 80.

The second cooling circuit 124 b comprises a supply passage 160, beingin fluid communication with the third inlet passage 92 and extendingfrom the root 82 to the tip 80. The supply passage 160 fluidlycommunicates with a trailing edge passage 162, extending from root 82 totip 80. In one example, the trailing edge passage 162 can be coupled tothe supply passage 160 through one or more inlets 164 extending radiallyalong the airfoil 78, disposed between the supply and trailing edgepassages 160, 162. The trailing edge passage 162 can be disposed withone or more rows of pins or pin banks 166 disposed within the trailingedge passage 162. The trailing edge passage 162 can further comprise oneor more slots 168 in fluid communication with the trailing edge passage162 and the exterior of the airfoil 78.

The mid cooling circuit 122 is disposed in the chord-wise middle of theairfoil 78 between the leading edge cooling circuit 120 and the trailingedge cooling circuit 124. The mid cooling circuit 122 can fluidly coupleto the second inlet passage 90, receiving a flow of cooling fluidtherefrom. The mid cooling circuit 122 comprises an upstream supplypassage 160, fluidly coupled to the second inlet passage 90, extendingfrom root 82 to tip 80. The upstream supply passage 160 can be locatedadjacent to the suction sidewall 100, or can be adjacent the pressuresidewall 98, or neither adjacent to the pressure nor suction sidewalls98, 100.

The upstream supply passage 170 fluidly couples to a mid supply passage172 extending from tip 80 to root 82. The mid supply passage 172 furthercouples to a mid return passage 174. A structural rib 190, spanning theinterior 96 and extending between the pressure sidewall 98 and thesuction sidewall 100, is disposed between the mid return passage 174 andthe combination of the upstream supply passage 170 and the mid supplypassage 172. The mid return passage 174 also can span the interior 96between the pressure sidewall 98 and the suction sidewall 100. Inaddition to the mid return passage 174, the mid supply passage 172further fluidly couples to a near wall cooling mesh 176 located adjacentto the pressure side between the pressure sidewall 98 and the mid supplypassage 172.

It should be appreciated that the geometry as illustrated in FIG. 4 isexemplary of one implementation of the cooling circuits disclosed hereinand should not be construed as limiting. The cooling circuits comprisinga plurality of passages, walls, channels, pin banks, etc. should beunderstood as one exemplary implementation of the cooling circuitswithin an airfoil 78, and the positions, dimensions, and geometriesdisclosed herein are incorporated in order to facilitate understandingof the inventive concept of the cooling circuits. For example, the nearwall cooling circuit 132, while shown on the suction sidewall 100 of theairfoil 78, can alternatively be located on the pressure sidewall 98.Additionally, the abstract shapes defined by the passages within thecross-sectional profile of the airfoil 78 are exemplary and can be anyshape, being geometrical, unique, or otherwise.

Looking at FIG. 5, best illustrating the mid cooling circuit 122, thenear wall cooling mesh 176 fluidly communicates with the upstream supplypassage 170 via the mid supply passage 172. The mid supply passage 172is in fluid communication with the near wall cooling mesh 176 through aradially extending opening, defining a fluid inlet 184 for the near wallcooling mesh 176. A dividing wall 186 is disposed between the mid supplypassage 172 and the near wall cooling mesh 176, partially defining themid supply passage 172, the inlet 184, and the near wall cooling mesh176. The dividing wall 186 terminates in a free end 188, adjacent to andpartially defining the inlet 184, located opposite of the structural rib190. The near wall cooling mesh 176 further comprises a channel 178having a plurality of pins or a pin bank 180 disposed within the channel178.

The channel 178 provides fluid communication between the mid supplypassage 172 and a mesh plenum 182. The mesh plenum 182 defines a fluidoutlet for the near wall cooling mesh 176. The mesh plenum 182 comprisesa geometry defining a cross-sectional area for the mesh plenum 182greater than the cross-sectional area of the near wall cooling mesh 176in the flow direction. In alternative orientations for the mid coolingcircuit 122, the near wall cooling mesh 176 can be located adjacent tothe suction sidewall 100.

It should be appreciated that the cross-sectional area for the meshplenum 182 is larger relative to the channel 178 facilitating machiningof a plurality of film holes along the span-wise length of the meshplenum 182. Typical drilling of film holes through the pressure orsuction sidewalls 98, 100 can extend into and damage additional walls orpassages within the interior 96 of the airfoil 78 if not provided withsufficient space for the drill to extend. The term “drill” as utilizedherein refers to machining by some known method, for example laser orelectro-discharge machining. The additional volume within the meshplenum 182, resultant from the increased cross-sectional area, can alsoeffect a kinder airflow within the mesh plenum 182, which can assist inestablishing a uniform cooling film exhausted from the film holes.

FIG. 6 is a flow diagram for the cooling circuits 120, 122, 124 of theairfoil 78 of FIG. 4. The airfoil 78 is schematically shown in brokenline to illustrate the general configuration of the cooling circuit 120,122, 124 within the airfoil 78. The airfoil 78 defines the interior 96as a cavity extending from the leading edge 102 to the trailing edge 104in a chord-wise direction and from the tip 80 to the root 82 in aspan-wise direction, and which can be divided into distinct channels orpassages by internal walls to form the cooling circuits 120, 122, 124,which direct a flow of cooling fluid through the airfoil 78. A tipcooling passage 208, disposed above the tip 80 of the airfoil 78, canextend in a substantially chord-wise direction from adjacent the leadingedge 102 toward the trailing edge 104. The tip cooling passage 208provides a common passage for the cooling circuits 120, 122, 124 toexhaust a cooling fluid, such that cooling fluid fed into the coolingcircuits 120, 122, 124 can be exhausted from the airfoil 78 if not beingexhausted through one or more film holes.

The leading edge cooling circuit 120 can be fed with cooling fluid fromthe first inlet passage 88 within the dovetail 76. The leading edgecooling circuit 120 receives the cooling fluid within the supply passage130 moving from the root 82 toward the tip 80. The supply passage 130fluidly communicates with the leading edge passage 144 through the leadcircuit channel 146 where a plurality of film holes 200 can exhaust thecooling fluid along the leading edge 102 of the airfoil 78 to create acooling film.

The leading edge cooling circuit 120 can further include at least oneupper turn 210 near the tip 80, providing cooling fluid to the near wallcooling circuit 132 from the supply passage 130. At the upper turn 210,the cooling fluid can flow from the supply passage 130 into plenumpassage 136. The cooling fluid travels within the plenum passage 136,comprising multiple pins 138, in a tip 80 to root 82 direction. Near theroot 82, the leading edge cooling circuit 120 can further comprise atleast one lower turn 212, exemplarily illustrated as two lower turns212, providing cooling fluid from the plenum passage 136 to the returnpassages 140. The cooling fluid flows within the return passages 140 ina root 82 to tip 80 direction, and can exhaust the cooling fluid throughthe film holes 200 to form a cooling film along the airfoil 78 exteriorsurface or move the cooling fluid toward the tip cooling passage 208.

The trailing edge cooling circuit 124 can be fed with a flow of coolingfluid from the third inlet passage 92. The third inlet passage 92, canfurther supply the first cooling circuit 124 a and the second coolingcircuit 124 b individually, which can be done by separating the coolingfluid flow from the third inlet passage 92 into a leading side inlet 92a and a trailing side inlet 92 b.

The first cooling circuit 124 a, illustrated as being fed from theleading side inlet 92 a, receives a cooling fluid flow within the supplypassage 150 in a root 82 to tip 80 direction. At an upper turn 218, thecooling fluid can be provided to the return passage 152 from the supplypassage 150, moving in a tip 80 to root 82 direction. The cooling fluidcan then be provided to the outlet passage 154 from the return passage152, at a lower turn 220 moving from root 82 to tip 80. Within theoutlet passage 154, the cooling fluid can be exhausted through the filmholes 200 to provide a cooling film along the exterior surface of theairfoil 78. Thus, the fluid path defined by the first cooling circuit124 a can be substantially serpentine, snaking between the tip 80 andthe root 82.

The second cooling circuit 124 b, illustrated as being fed from thetrailing side inlet 92 b, can be provided with a flow of cooling fluidin the supply passage 160 in a root 82 to tip 80 direction. Along thesupply passage 160, cooling fluid can flow into a trailing edge passage162, which can comprise one or more pins 166, through one or more inlets164 from the supply passage 160. Cooling fluid, which does not flow intothe trailing edge passage 162, can be exhausted from the tip end 222 ofthe supply passage out the trailing edge 104 of the airfoil 78. Coolingair within the trailing edge passage 162 can be exhausted from theairfoil 78 through the film holes 200, or can alternatively be exhaustedthrough the trailing edge 104 through the slots 168.

The mid cooling circuit 122, disposed chord-wise between the leadingedge cooling circuit 120 and the trailing edge cooling circuit 124, canbe fed with a flow of cooling fluid from the second inlet passage 90.The mid cooling circuit 122 receives the cooling fluid within theupstream supply passage 170 from the second inlet passage 90, movingfrom the root 82 toward the tip 80. The mid cooling circuit 122 furthercomprises an upper turn 214 where the upstream supply passage 170fluidly communicates with the mid supply passage 172. The mid supplypassage 172 is further in fluid communication with the near wall coolingmesh 176 through the inlet 184. The cooling airflow from the inlet 184can move through the near wall cooling mesh 176, traveling from theinlet 184 and through the channel 178, which can include one or morepins or pin banks 180, to the mesh plenum 182 where the cooling fluidcan be exhausted through the film holes 200 to create a cooling filmalong the external surface of the airfoil 78.

The mid cooling circuit 122 can further comprise a lower turn 216,providing cooling fluid to the mid return passage 174 from the midsupply passage 172. From the mid return passage 174, the film holes 200can exhaust the cooling fluid from the mid cooling passage 122 toprovide a cooling film along the exterior surface of the airfoil 78. Themid return passage 174 extends between the suction and pressuresidewalls 98, 100, such that sets of film holes 200 can provide acooling fluid flow to the exterior surfaces of the airfoil 78.

It should be appreciated that cooling circuits 120, 122, 124, asillustrated in FIG. 6 are exemplary of one implementation of the coolingcircuits within an airfoil 78 and should not be construed as limited bythe particular geometry, passages, pin banks, film holes, etc. It shouldbe further understood that while the cooling circuits 120, 122, 124, 124a, 124 b are illustrated as generally moving from the leading edge 102toward the trailing edge 104 or the trailing edge 104 toward the leadingedge 102, the illustration is only an exemplary depiction of the coolingcircuits themselves. The particular passages, channels, inlets, or meshcan flow in any direction relative to the airfoil 78, such as in thetrailing or leading edge 102, 104 direction, tip 80 or root 82direction, or toward the pressure 98 or suction 100 sidewalls of theairfoil 78, or any combination thereof.

FIG. 7 illustrates a cross-section of an airfoil 278 where a near wallcooling mesh 376 within a mid cooling circuit 322 extends in chord-wisedirection being substantially of the trailing edge 104 toward theleading edge 102. The airfoil 278 of FIG. 7 can be substantially similarthe airfoil 78 of FIG. 4. As such, similar elements will be identifiedwith similar numerals increased by a value of two-hundred.

The mid cooling circuit 322 is fed from the second inlet passage 90,being in fluid communication with an upstream supply passage 370. Theupstream supply passage 370, extending in a root 82 to tip 80 direction,is in fluid communication with a mid supply passage 372. The mid supplypassage 372 extends from tip 80 to root 82, and is in further fluidcommunication with the near wall cooling mesh 376 and the mid returnpassage 374. The mid return passage 374 receives a cooling fluid formthe mid supply passage 372 near the root 82, and extends from root 82 totip 80. Film holes can extend from the external surface of the airfoil78 to the mid return passage 374, providing the cooling fluid to theexternal surface of the airfoil 278 to form a cooling film.

The near wall cooling mesh 376 comprises a channel 378, being fed fromthe mid supply passage 372 through a radially extending opening defininga fluid inlet 384. The channel 378 in FIG. 7, in contrast to the channel178 of FIG. 4, extends in a trailing edge 104 to leading edge 102chord-wise direction. A plurality of pins or pin banks 380 can bedisposed within the channel 378. A mesh plenum 382 is in fluidcommunication with the mid supply passage 372 via the channel 378 and isdisposed chord-wise-opposite of the inlet 384 relative to the channel378. A dividing wall 388 further defines and is disposed between the midsupply passage 372 and the channel 378. The dividing wall 388 comprisesa free end 388 adjacent to the inlet 384.

The mesh plenum 382 can be disposed adjacent to the mid return passage374, such that the mesh plenum 382 is situated between the mid returnpassage 374 and the pressure sidewall 98. A plurality of film holes canextend into the mesh plenum 382, providing the cooling fluid to theexternal surface of the airfoil 278 to form a cooling film. It should beappreciated that while the near wall cooling mesh 376 is illustrated asadjacent to the pressure sidewall 98, it can alternatively be disposedadjacent to the suction sidewall 100.

The mesh plenum 382 comprises a cross-sectional area being largerrelative to the channel 378, facilitating machining of the film holesextending between the external surface of the airfoil and the meshplenum 382. Typical drilling of film holes through the sidewalls 98,100, can extend into and damage interior walls or passages if notprovided with sufficient space for the drill to extend. The additionalvolume within the mesh plenum 382 further provides a kinder airflowwithin the mesh plenum 382, which can assist in establishing a uniformcooling film exhausted form the film holes.

It should be appreciated that the geometry as illustrated in FIG. 7 canbe exemplary of one implementation of the mid cooling circuit 322, andthat particular geometries of the passages, wall, channels, pin banks,etc., can vary from the airfoil 278 as shown. For example, thecross-sectional area of the combination of the mid supply passage 372and the near wall cooling mesh 376 can be smaller or disposed closer tothe trailing edge 104, such that the mid return passage 374 spans theinterior between the pressure sidewall 98 and the suction sidewall 100.

It should be further appreciated that mid cooling circuits 122, 322 asillustrated in FIGS. 4, 5 and 7 are exemplary implementations of thecooling circuits comprising a near wall cooling mesh 176, 376 and a meshplenum 182, 382 within an airfoil 78, 278 and should not be construed aslimited by the orientation of the passages, pin banks, film holes, etc.It should be further understood that while the cooling circuits 120,122, 124, 124 a, 124 b are illustrated as generally moving from theleading edge 102 toward the trailing edge 104 or the trailing edge 104toward the leading edge 102, the illustration is only an exemplarydepiction of the cooling circuits themselves. The particular passages,channels, inlets, or mesh can flow in any direction relative to theairfoil 78, such as in the trailing or leading edge 102, 104 direction,tip 80 or root 82 direction, or toward the pressure or suction sidewalls98, 100 of the airfoil 78, or any combination thereof.

The various embodiments of systems, methods, and other devices disclosedherein provide improved cooling effectiveness for the cooling circuit ofa turbine blade. One advantage that may be realized in the practice ofsome embodiments of the described systems is that the near wall coolingmesh of the blade can be utilized with at least one of pressure orsuction sidewalls of the blade, or with both sidewalls, while providingair to the film holes in order to create a cooling film on the externalsurface of the airfoil. The implemented near wall cooling mesh providesoptimal cooling and airflow management within an airfoil, and providesspace for the machining of film holes within the near wall cooling mesh.While the specific embodiments are described in terms of an airfoil inthe form of a turbine blade, the description is equally applicable toany airfoil within the gas turbine engine, including, withoutlimitation, turbine vanes, compressor blades and compressor vanes.

This written description uses examples to disclose the invention,including the best mode, and to enable any person skilled in the art topractice the invention, including making and using any devices orsystems and performing any incorporated methods. The patentable scope ofthe invention is defined by the claims, and can include other examplesthat occur to those skilled in the art. Such other examples are intendedto be within the scope of the claims if they have structural elementsthat do not differ from the literal language of the claims, or if theyinclude equivalent structural elements with insubstantial differencesfrom the literal languages of the claims.

What is claimed is:
 1. An airfoil for a gas turbine engine, the airfoilcomprising: an outer surface defining a pressure side and a suction sideextending axially between a leading edge and a trailing edge andextending radially between a root and a tip; and a cooling circuitlocated within the airfoil and comprising: a cooling air inlet passage;a radially extending supply passage; an upstream supply passage fluidlycoupling the cooling air inlet passage and the supply passage; a nearwall cooling mesh located adjacent to and extending along a portion ofthe outer surface; and a radially extending opening fluidly coupling thesupply passage to the near wall cooling mesh to define a fluid inlet forthe near wall cooling mesh; wherein cooling air flows in a firstdirection through the upstream supply passage and in a second directionthrough the supply passage, with the second direction being atip-to-root direction, and wherein the cooling air flows through thenear wall cooling mesh via the fluid inlet.
 2. The airfoil according toclaim 1 further comprising a pin bank located within the near wallcooling mesh.
 3. The airfoil according to claim 1 wherein the near wallcooling mesh is located adjacent to one of the pressure side and thesuction side.
 4. The airfoil according to claim 1 further comprising aradially extending dividing wall located within an interior of theairfoil and extending along the outer surface to partially define thesupply passage and the near wall cooling mesh.
 5. The airfoil accordingto claim 4 wherein the dividing wall terminates in a free end partiallydefining the fluid inlet for the near wall cooling mesh.
 6. The airfoilaccording to claim 5 further comprising a structural rib located withinthe interior and extending between the pressure side and the suctionside, with the structural rib partially defining the supply passage andspaced from the free end to define the fluid inlet in combination withthe free end.
 7. The airfoil according to claim 1, further comprising aplenum fluidly coupled to the near wall cooling mesh to define a fluidoutlet for the near wall cooling mesh.
 8. The airfoil according to claim7 wherein one of the fluid inlet and the fluid outlet for the near wallcooling mesh is located closer to the leading edge than the other of thefluid inlet and the fluid outlet for the near wall cooling mesh.
 9. Theairfoil according to claim 7 wherein a maximum cross-sectional area ofthe plenum is greater than a maximum cross-sectional dimension of thenear wall cooling mesh.
 10. The airfoil according to claim 7 furthercomprising a film hole extending through the outer surface and into theplenum.
 11. The airfoil according to claim 7 wherein the near wallcooling mesh is arranged such that the cooling air flowing through nearwall cooling mesh from the fluid inlet to the fluid outlet flows in adirection from the leading edge toward the trailing edge.
 12. Theairfoil according to claim 7 wherein at least a portion of the plenumextends along the outer surface.
 13. The airfoil according to claim 7wherein a cross-sectional area of the plenum is greater than across-sectional area of the near wall cooling mesh in an airflowdirection.
 14. A blade for a gas turbine engine having a turbine rotordisk, the blade comprising: a dovetail having multiple air inletpassages and configured to mount to the turbine rotor disk; an airfoilextending radially from the dovetail and having an outer surfacedefining a pressure side and a suction side extending axially between aleading edge and a trailing edge and extending radially between a rootand a tip, with the root being adjacent the dovetail; a leading edgecooling circuit located within the airfoil and fluidly coupled to afirst air inlet passage of the multiple air inlet passages; a trailingedge cooling circuit located within the airfoil and fluidly coupled to asecond air inlet passage of the multiple air inlet passages; and a midcooling circuit located within the airfoil and between the leading edgecooling circuit and the trailing edge cooling circuit, the mid coolingcircuit comprising: a radially extending supply passage; an upstreamsupply passage fluidly coupling a third air inlet passage of themultiple air inlet passages to the supply passage; a near wall coolingmesh located adjacent to and extending along a portion of the outersurface; and a radially extending opening fluidly coupling the supplypassage to the near wall cooling mesh to define a fluid inlet for thenear wall cooling mesh; wherein cooling air flows in a first directionthrough the upstream supply passage and in a second direction throughthe supply passage, with the second direction being a tip-to-rootdirection, and wherein the cooling air flows through the near wallcooling mesh via the fluid inlet.
 15. The blade according to claim 14further comprising a radially extending dividing wall located within aninterior of the airfoil and extending along the outer surface topartially define the supply passage and the near wall cooling mesh. 16.The blade according to claim 15 wherein the dividing wall terminates ina free end partially defining the fluid inlet for the near wall coolingmesh.
 17. The blade according to claim 16 further comprising astructural rib located within the interior and extending between thepressure side and the suction side, with the structural rib partiallydefining the supply passage and spaced from the free end to define thefluid inlet in combination with the free end.
 18. The blade according toclaim 14, further comprising a plenum fluidly coupled to the near wallcooling mesh to define a fluid outlet for the near wall cooling mesh,wherein a cross-sectional area of the plenum is greater than across-sectional area of the near wall cooling mesh in an airflowdirection.
 19. The blade according to claim 18 wherein one of the fluidinlet and the fluid outlet for the near wall cooling mesh is locatedcloser to the leading edge than the other of the fluid inlet and thefluid outlet for the near wall cooling mesh.
 20. An airfoil for a gasturbine engine having a peripheral wall bounding an interior anddefining a pressure side and a suction side, opposite the pressure side,a near wall cooling mesh located adjacent to and extending along aportion of one of the suction side and pressure side, a supply passage,a cooling air inlet passage, and an upstream supply passage fluidlycoupling the cooling air inlet passage and the supply passage, whereincooling air flows in a first direction through the upstream supplypassage and a second direction through the supply passage, with thesecond direction being a tip-to-root direction, and wherein the coolingair flows through the near wall cooling mesh via the cooling air inletpassage.